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Introduction
Introduction
The thermal control subsystem is concerned with monitoring and maintaining temperatures across the satellite within their allowable limits.
Design Overview
Design Overview
The main functions of the thermal control subsystem are to:
Maintain all satellite components within their allowable temperature ranges while spacecraft is in the expected operating conditions (R-MIS-0110)
Monitor spacecraft temperatures as necessary for health monitoring, heater control, and necessary science data collection
These functions are outlined in the following functional block diagram.
The bill of materials for the primary heater is under a non disclosure agreement, it is currently under review by NASA.
The redundant Minco heater is certified by both NASA and the ESA.
System Architecture Diagram
System Architecture Diagram
The following diagram shows the architecture of the thermal control subsystem, and how all the components are connected to each other and to the rest of the spacecraft, both physically and electrically.
The thermal control components used for the satellite will be mainly passive, with the exception of heaters. The thermal control subsystem will largely consist of using different materials and coatings to control the radiation properties and flow of heat through the satellite. All exterior structure surfaces that are made of aluminum will be black anodized to maximize heat absorbed, as the satellite body runs cold. The solar panels typically absorb much more heat than the spacecraft structure, and due to their low mass and thickness can heat up very quickly. Therefore, all solar panels will be made using a white solder mask to minimize absorbed heat. A conductive path from the solar panels to the main spacecraft body will be formed by the deployment hinges.
Two different thermistors are used in the thermal subsystem- one part mount thermistor with wire leads, and one board mount thermistor that is soldered directly to the mounting location. The part mount thermistor is used in only one location, on the payload sample plate. All other thermistors are located on electronic boards where they can be soldered in place. The part mount thermistor will be mounted using thermal epoxy (Stycast 2850 with catalyst 9) and then wired in to the payload control board. All other thermistors will be soldered to their appropriate mounting locations and then staked in place using thermal epoxy (Stycast 2850 with catalyst 9).
Due to the requirement for 80% system reliability (R-MIS-0003 ), we have decided to include redundant heaters on the batteries. This is based on our decision to use Datec 3D printed heaters as our primary battery heaters. These heaters do not have space flight heritage, and therefore it is difficult to prove that they will meet the reliability requirement in a space environment. Space-proven Minco heaters will be included in the design to provide redundant heating capabilities, and meet the 80% reliability requirement. A redundant battery thermistor is also included in the thermal subsystem design as accurate battery temperature readings are important to battery health. If both thermistors are operating properly, the average temperature of the two will be used to control the heater power. If it is clear that one thermistor has failed, the data from the failed thermistor will be neglected and only one thermistor will be used for calculations.
The redundant heater will only be used if the primary heater fails, or if the spacecraft enters a low power mode where the C&DH is powered down. The power to the primary heater is controlled by a switch, while the power to the redundant heater is not switched. The flow of power to the redundant heater will be controlled by a mechanical thermostat wired in line with the power line from the PCU.
During typical operating modes (any mode where the state of charge of the battery is above 40%) , the heating of the batteries is controlled by the C&DH, which reads the temperatures from the battery thermistors and determines if the heater should be turned on or off. The control algorithm (temperature cutoffs) for the primary battery heater will be defined such that the thermostat will be creating an open circuit (no power flowing) at all times when the PCU is on, to ensure that both heaters are not activated at the same time.
When the state of charge of the batteries falls below 40%, the C&DH is turned off to conserve power. When the C&DH is powered down, the PCU will set the switch so that the primary (Datec) heater is off. At this point the redundant battery heater will be relied on to control the battery temperature, with the mechanical thermostat opening and closing as the battery temperature varies in order to power the heater on and off. The mechanical thermostat will be mounted on the battery saddle, in close proximity to the batteries. The function flow for the thermostat operation is shown in the diagram below.
In the case where the primary heater has failed, as determined by telemetry, the switch to the primary heater will be turned off permanently and the redundant heaters with thermostat control will be relied on entirely to maintain battery temperatures. This process will also occur automatically if the primary heater switch fails open (no flow of power). A switch that has a higher likelihood of failing open will be used here, to minimize the chance of a fail closed resulting in constant power to the battery heater.
Thermistor Locations
Thermistor Locations
The following table describes the locations where thermistors will/should be located on the spacecraft. Ranking indicates how important it is to have the data: 1 means data is necessary either for spacecraft health or for payload data, a 2 means the data is highly desired and could be used for health monitoring but is not required, and a 3 means data may be useful but is not of high priority. Thermistors for health monitoring would have their readings taken most frequently, and sent to ground with every pass. Other thermistors would be logged and sent less frequently.
Design Trade-Offs
Design Trade-Offs
RTDs vs. Thermistors
RTDs vs. Thermistors
We chose thermistors for temperature measurement on board Iris for a number of reasons. First of all, thermistors are generally cheaper than RTDs when expected temperatures are within the standard measurement limitations of thermistors. Since the expected temperatures experienced by Iris are fairly mild, we are well within the temperature range of standard automotive-grade thermistors. Thermistors also have a faster response time and a smaller volume when compared to RTDs, as they are often available with a pin head sized temperature sensing bulb, which has an extremely low thermal capacitance compared to larger RTDs. The one potential advantage of RTDs over thermistors for Iris is that RTDs have a linear relationship between temperature and resistance, while thermistor resistance response is non-linear. However, equations that define the resistance profile of thermistors are provided by manufacturers, and the non-linear behaviour can therefore be accounted for by the software. For these reasons, thermistors are chosen as the preferable temperature sensor.
Initial Analysis
Initial Analysis
MatLab Analysis
MatLab Analysis
For the thermal control of Manitoba-SAT1, it is expected that passive thermal control techniques for cooling and active control for heating will be sufficient to maintain components within their allowable temperature ranges. This assumption is based on a simplified nodal analysis of the spacecraft in Matlab which placed a node in the center of each spacecraft panel and considered all heat flows in and out of that surface. This analysis calculated spacecraft minimum and maximum temperatures over a number of orbits of approximately -15C and +60C when the spacecraft is operating with no thermal control subsystem. The sawtooth pattern that is seen in the graph is due to the cyclical external heat loads coming from the spacecraft moving in and out of orbit. As the data shows, after a few orbits the spacecraft stabilizes to a consistent cyclical temperature profile. Comparing these temperatures to the allowable operating temperatures of all spacecraft units and components, many of which are automotive grade, use of coatings, radiators, and heaters should be sufficient to control the spacecraft temperatures to the desired level.
Results of MatLab nodal analysis of orbiting spacecraft is shown above (blue line shows lower temperature limit of automotive grade components [-30C] , red line shows upper temperature limit [+85C]).
Phase B Orbital Simulations
Phase B Orbital Simulations
Temperatures reached during the thirty minute hold period following ejection from the NRCSD were a concern for the thermal control subsystem, and specifically for the batteries as they have the strictest temperature limits, as no heaters can be operated during this period. Simulations were performed to determine the worst case cold temperature that will be reached by the batteries during the hold. For these simulations, we assumed the satellite had the high emissivity of black anodized aluminum and was in eclipse for the entire thirty minute period. The solar panels were stowed, and connected to the satellite structure with conductive couplings in the hinge locations. The long axis of the satellite was aligned with the orbit normal, and the satellite was rotating about the long axis, with an extremely low angular rotation rate such that one of the non-solar panel sides was facing deep space for the entire simulation period . The initial temperature of the satellite was set to -10C, which is the worst case pre-deployment temperature as defined by NanoRacks.
The images to the left show the temperature of the cubesat at the end of thirty minute hold. The image on the left shows the satellite with the solar panels, and the image on the right is without the solar panels. The battery is located in the top module of the spacecraft, as seen in these images.
The images below shows the battery with and without the surrounding module. As we can see, the thermal mass of the battery is large enough that the battery temperature is nearly unaffected by the surrounding environment and changes by only about a degree during the hold period.
Note that these simulations are also used as verification of requirement R-THE-0780, requiring that the thermal subsystem be able to recover from a 5 minute power upset. The only thermal concern caused by a power upset would be the batteries getting too cold. As these simulations show, even during a 30 minute period the batteries will only cool down by approximately 1.5C worst case. Considering this, a 5 minute period with no power should have a negligible impact on battery temperatures, even when the spacecraft is in eclipse.
Key Risks
Key Risks
Phase B
Phase B
Risk D42 was a key concern for the Phase B design, which has been mitigated by the orbital simulations shown above.
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